The present invention relates to the general field of combustion chambers for turbomachines, whether for aviation or terrestrial purposes.
Typically, an aviation or terrestrial turbomachine is made up of an assembly that comprises in particular an annular compression section for compressing the air passing through the turbomachine, an annular combustion section disposed at the outlet from the compression section and within which the air coming from the compression section is mixed with fuel in order to be burnt therein, and an annular turbine section disposed at the outlet from the combustion section and having a rotor that is driven in rotation by the gas coming from the combustion section.
The compression section is in the form of a plurality of stages of moving wheels, each carrying blades that are placed in an annular channel through which the turbomachine air passes, the channel being of section that decreases going from upstream to downstream. The combustion section comprises a combustion chamber in the form of an annular channel in which the compressed air is mixed with fuel in order to be burnt therein. The turbine section is made up of a plurality of stages of moving wheels, each carrying blades that are placed in an annular channel through which the combustion gas passes.
The flow of air through the above assembly generally takes place as follows: the compressed air coming from the last stage of the compression section presents natural gyratory motion with an angle of inclination of the order of 35° to 45° relative to the longitudinal axis of the turbomachine, which angle of inclination varies as a function of the speed of rotation of the turbomachine. On entry into the combustion section, this flow of compressed air is straightened along the longitudinal axis of the turbomachine (i.e. the angle of inclination of the air relative to the longitudinal axis of the turbomachine is reduced to 0°) by means of a guide vane. The air in the combustion chamber is then mixed with the fuel so as to provide satisfactory combustion, and the gas generated by the combustion continues to flow generally along the longitudinal axis of the turbomachine in order to reach the turbine section. Once there, the combustion gas is redirected by a nozzle in order to present gyratory motion with an angle of inclination greater than 70° relative to the longitudinal axis of the turbomachine. Such an angle of inclination is essential for producing the angle of attack that is needed to provide the mechanical force for driving rotation of the moving wheel of the first stage of the turbine section.
Such a distribution of angles of inclination for the air passing through the turbomachine presents numerous drawbacks. The air that naturally leaves the last stage of the compression section with an angle of inclination lying in the range 35° to 45° is successively straightened (its angle is reduced to 0°) on entry into the combustion section, and then redirected to have an angle of inclination greater than 70° on entry into the turbine section. These successive changes of angle of inclination in the flow of air through the turbomachine require intense aerodynamic forces to be produced both by the guide vane of the compression section and by the nozzle of the turbine section, which aerodynamic forces are particularly damaging to the overall efficiency of the turbomachine.